Gas turbine engines and similar turbomachines have one or more compressors and turbines. The compressors and turbines include longitudinally alternating arrays of blades and vanes that extend radially across an annular flowpath. During operation of the turbomachine, a working medium fluid flows longitudinally through the flowpath. The blades and vanes interact with the working medium fluid to transfer energy from the compressor to the fluid and from the fluid to the turbine.
A typical turbine blade has a root that adapts the blade to be secured to a rotatable hub so that the blade extends radially outwardly from the hub. The blade also includes a platform adjacent to the root, a shroud radially spaced apart from the platform and an airfoil spanning between the platform and the shroud. The airfoil has an intangible, spanwisely extending stacking line, which is a manufacturing reference for establishing the contour of the airfoil and the spatial relationship of the airfoil relative to the platform and shroud. The blade also has a knife edge that extends outwardly from the radially outer surface of the shroud. When a full complement of blades is secured to the hub of a turbomachine to form a blade array, the blade platforms and shrouds define radially inner and outer boundaries of a flowpath for the working medium fluid, and the airfoils span radially across the flowpath. In addition, the knife edges of the installed blades abut each other to form a substantially circumferentially continuous knife edge ring. The knife edge ring extends radially toward an abradable seal that circumscribes the blade array. Over the course of a brief break-in period early in the life of a newly manufactured turbomachine (or after an engine has is been refurbished with a new abradable seal) mechanical deflections and thermally induced dimensional changes cause the knife edge ring to cut a corresponding knife edge groove into the abradable seal. Thereafter, the engine operates with the knife edge ring protruding snugly into the knife edge groove thereby forming a seal that minimizes the leakage of working medium fluid past the airfoils.
The turbine blades are made from a high strength, temperature tolerant alloy that is cast to a near net shape. The blade casting is then polished to smooth out any minor irregularities in the airfoil surface and to remove any excess material from the airfoil's leading and trailing edges. The blade platform and shroud are then finish machined to render the blade dimensionally and geometrically suitable for installation and service in the turbomachine.
The finish machining of the platform and shroud is conducted according to a highly automated manufacturing protocol known as "one piece flow". This protocol features a sequence of machines arranged in a logistically optimized flow line in a manufacturing facility. Each machine performs one of several required machining steps, and operates on only one blade at a time. Each blade is transferred from machine to machine in the flow line. In order to guard against the accumulation of machining inaccuracies due to these transfers, the blade has a set of dedicated, highly accurate machining datums.
One of the dedicated datums found on conventional blades is a projection that extends radially outwardly from the shroud. The projection has a radially inner portion that is recessed in a pocket in the shroud outer surface and blends into a pocket sidewall so that the radially inner portion of the projection is peripherally noncontinuous. The projection also has a radially outer portion that extends radially beyond the sidewall and therefore is peripherally continuous. The peripherally continuous outer portion serves not only as a reference point but also as one of a number of interfaces by means of which the blade can be predictably positioned and anchored in place in each of the one piece flow machines.
If such a blade were to be installed in a turbomachine, the outer portion of the projection would cut into the seal that circumscribes the blade array in much the same way that the knife edge ring cuts into the seal. Because the knife edge ring and the projection are separated by only a small distance, and because the hub shifts slightly in the longitudinal direction during engine operation, the cut made by the projection can merge with and expand the width of the knife edge groove. As a result the desired snug fit between the knife edge ring and the knife edge groove is undermined thereby diminishing the effectiveness of the seal and reducing the turbomachine's efficiency. To guard against this occurrence, the outer portion of the projection is machined off, leaving behind only the inner portion, which is too short to contact the seal. The removal of the outer portion also destroys the projection's utility as a datum, and therefore the outer portion is not removed until all the other machining operations are complete and the projection is no longer necessary to facilitate blade manufacture.
Turbine blades, such as those just described, must also undergo post-manufacturing dimensional inspections. The above described machining projection would be an ideal inspection reference point. However since the projection is not available as a reference in the post-manufacturing environment, it is common practice to rely on alternative reference points. These alternative references are two sets of datum triplets on the airfoil, one set near the platform and one near the shroud. These datum triplets are far less accurate than the machining projection, and therefore overly stringent inspection criteria must be established to ensure that all dimensionally unacceptable blades are identified. As a result, some blades that are dimensionally acceptable will be identified as unacceptable, and will be scrapped. Given the high cost of turbomachinery blades it is clearly desirable to minimize the likelihood that a serviceable blade is erroneously identified as unacceptable.
A similar difficulty arises when damaged or deteriorated blades are refurbished to extend their serviceability. The absence of the machining projection forces reliance upon less accurate datums, and as a result it may not be possible to realize the full potential of the refurbishment.
One way to ensure that the projection's utility survives the original manufacturing process is described in commonly owned, copending U.S. patent application 08/953129, entitled "Turbomachinery Blade or Vane with a Permanent Machining Datum", filed on Oct. 17, 1997. The '129 application discloses a turbine blade having a conical datum projecting from a pocket in the blade shroud and spaced a distance s from the pocket sidewall. Because of the spacing, the peripheral continuity of the projection survives the finish machining operations so that the projection retains its utility as a reference and anchoring feature. Because the projection is partially recessed within the pocket, it is partially protected from damage due to careless handling. In addition, the axis of the projecting datum is aligned with (i.e. coincides with) the airfoil stacking line to simplify manufacturing tooling requirements.
The survivable datum disclosed in the above noted patent application is superior to a nonsurvivable datum, but is not without certain shortcomings. Although much of the projection is recessed within a pocket, the tip of the projection is exposed and therefore vulnerable to possible handling damage. Moreover, in the disclosed blade, the highly desirable alignment of the datum axis with the stacking line is obtained by introducing a scallop into the pocket sidewall. In blades that require a generous scallop, the presence of the scallop can degrade the stiffness of the shroud, making it susceptible to undesirable, centrifugally induced deflections during engine operation. Although a generous scallop could be avoided by repositioning the datum elsewhere on the shroud, such repositioning would separate the datum axis from the stacking line, thereby complicating the tooling required for blade manufacture.